Airfoil Comparison
Select airfoils from the airfoil database or add your own airfoils and compare the airfoil shape and lift/drag polars.
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(naca66209-il) NACA 66-209 | NACA 66-209 airfoil Max thickness 9% at 45% chord Max camber 1.1% at 50% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca66209-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca66209-il | 50,000 | 9 | 22.1 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca66209-il | 50,000 | 5 | 25.9 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca66209-il | 100,000 | 9 | 32.4 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca66209-il | 100,000 | 5 | 40.6 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca66209-il | 200,000 | 9 | 53.3 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca66209-il | 200,000 | 5 | 47.1 at α=3° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca66209-il | 500,000 | 9 | 67.7 at α=2.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca66209-il | 500,000 | 5 | 60.6 at α=2° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca66209-il | 1,000,000 | 9 | 80.4 at α=2° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca66209-il | 1,000,000 | 5 | 64.9 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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